Gas turbine engine airfoil

ABSTRACT

A compressor airfoil of a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between leading edge sweep angle and span position defined by a curve in which the leading edge sweep angle is positive at 0% span and crosses to a negative leading edge sweep angle at a span position less than 80% span. A negative sweep angle is in the forward direction. A positive sweep angle is in the rearward direction. The airfoil has a relationship between trailing edge sweep angle and span position defined by a curve in which the trailing edge sweep angle is negative at 0% span and changes less than 10° from 0% span to 90% span.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation to U.S. application Ser. No.15/114,745, filed on Jul. 27, 2016, which is a United States Nationalphase of International Application No. PCT/US2015/016091, filed on Feb.17, 2015, which claims priority to U.S. Provisional Application No.61/941,707 filed on Feb. 19, 2014.

BACKGROUND

This disclosure relates to gas turbine engine airfoils. Moreparticularly, the disclosure relates to airfoil leading and trailingedge sweep in, for example, a gas turbine engine compressor.

A turbine engine such as a gas turbine engine typically includes a fansection, a compressor section, a combustor section and a turbinesection. Air entering the compressor section is compressed and deliveredinto the combustor section where it is mixed with fuel and ignited togenerate a high-speed exhaust gas flow. The high-speed exhaust gas flowexpands through the turbine section to drive the compressor and the fansection. The compressor section typically includes at least low and highpressure compressors, and the turbine section includes at least low andhigh pressure turbines.

Direct drive gas turbine engines include a fan section that is drivendirectly by one of the turbine shafts. Rotor blades in the fan sectionand a low pressure compressor of the compressor section of direct driveengines rotate in the same direction.

Gas turbine engines have been proposed in which a geared architecture isarranged between the fan section and at least some turbines in theturbine section. The geared architecture enables the associatedcompressor of the compressor section to be driven at much higherrotational speeds, improving overall efficiency of the engine. Thepropulsive efficiency of a gas turbine engine depends on many differentfactors, such as the design of the engine and the resulting performancedebits on the fan that propels the engine and the compressor sectiondownstream from the fan. Physical interaction between the fan and theair causes downstream turbulence and further losses. Although some basicprinciples behind such losses are understood, identifying and changingappropriate design factors to reduce such losses for a given enginearchitecture has proven to be a complex and elusive task.

Prior compressor airfoil geometries may not be suitable for thecompressor section of gas turbine engines using a geared architecture,since the significantly different speeds of the compressor changes thedesired aerodynamics of the airfoils within the compressor section.Counter-rotating fan and compressor blades, which may be used in gearedarchitecture engines, also present design challenges.

SUMMARY

In one exemplary embodiment, a compressor airfoil of a turbine enginehaving a geared architecture includes pressure and suction sidesextending in a radial direction from a 0% span position to a 100% spanposition. The airfoil has a relationship between leading edge sweepangle and span position defined by a curve in which the leading edgesweep angle is positive at 0% span and crosses to a negative leadingedge sweep angle at a span position less than 80% span. A negative sweepangle is in the forward direction, and a positive sweep angle is in therearward direction.

In a further embodiment of the above airfoil, the curve crosses to thenegative leading edge sweep angle at a span position less than 75% span.

In a further embodiment of any of the above airfoils, the curve crossesto the negative leading edge sweep angle in a range of 60% span to lessthan 75% span.

In a further embodiment of any of the above airfoils, the curve crossesto the negative leading edge sweep angle at a span position less than40% span.

In a further embodiment of any of the above airfoils, the airfoil has arelationship between trailing edge sweep angle and span position definedby a curve in which the trailing edge sweep angle is negative at 0% spanand changes less than 10° from 0% span to 90% span.

In a further embodiment of any of the above airfoils, the airfoil has arelationship between trailing edge sweep angle and span position definedby a curve in which the trailing edge sweep angle is positive at 0% spanand has a decrease in trailing edge sweep angle of greater than 10° from0% span to 40% span.

In a further embodiment of any of the above airfoils, the positive sweepangle is in the rearward direction with respect to the orientationestablished by the velocity vector.

In one exemplary embodiment, a gas turbine engine includes a combustorsection arranged between a compressor section and a turbine section. Afan section has an array of twenty-six or fewer fan blades. The fansection has a low fan pressure ratio of less than 1.55. A gearedarchitecture couples the fan section to the turbine section or thecompressor section. An airfoil arranged in the compressor sectionincludes pressure and suction sides extending in a radial direction froma 0% span position at an inner flow path location to a 100% spanposition at an airfoil tip. The airfoil has a relationship betweenleading edge sweep angle and span position defined by a curve in whichthe leading edge sweep angle is positive at 0% span and crosses to anegative leading edge sweep angle at a span position less than 80% span.A negative sweep angle is in the forward direction, and a positive sweepangle is in the rearward direction.

In a further embodiment of the above gas turbine engine, the compressorsection includes at least a low pressure compressor and a high pressurecompressor. The high pressure compressor is arranged immediatelyupstream of the combustor section.

In a further embodiment of any of the above gas turbine engines, theairfoil is provided in a compressor outside the high pressurecompressor.

In a further embodiment of any of the above gas turbine engines, the lowpressure compressor is counter-rotating relative to the fan blades.

In a further embodiment of any of the above gas turbine engines, the gasturbine engine is a two-spool configuration.

In a further embodiment of any of the above gas turbine engines, the lowpressure compressor is immediately downstream from the fan section.

In a further embodiment of any of the above gas turbine engines, theairfoil is rotatable relative to an engine axis of an engine staticstructure.

In a further embodiment of any of the above gas turbine engines, thecurve crosses to the negative leading edge sweep angle at a spanposition less than 75% span.

In a further embodiment of any of the above gas turbine engines, thecurve crosses to the negative leading edge sweep angle in a range of 60%span to less than 75% span.

In a further embodiment of any of the above gas turbine engines, thecurve crosses to the negative leading edge sweep angle at a spanposition less than 40% span.

In a further embodiment of any of the above gas turbine engines, theairfoil has a relationship between trailing edge sweep angle and spanposition defined by a curve in which the trailing edge sweep angle isnegative at 0% span and changes less than 10° from 0% span to 90% span.

In a further embodiment of any of the above gas turbine engines, theairfoil has a relationship between trailing edge sweep angle and spanposition defined by a curve in which the trailing edge sweep angle ispositive at 0% span and has a decrease in trailing edge sweep angle ofgreater than 10° from 0% span to 40% span.

In a further embodiment of any of the above gas turbine engines, thepositive sweep angle is in the rearward direction with respect to theorientation established by the velocity vector.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment with ageared architecture.

FIG. 2 schematically illustrates a low pressure compressor section ofthe gas turbine engine of FIG. 1.

FIG. 3 is a schematic view of airfoil span positions.

FIG. 4 is a schematic view of a cross-section of an airfoil sectioned ata particular span position and depicting directional indicators andvelocity vectors in relation to the leading and trailing edges.

FIG. 5 is a schematic perspective view of an airfoil fragmentillustrating the definition of a leading edge sweep angle.

FIG. 6 is a schematic perspective view of an airfoil fragmentillustrating the definition of a trailing edge sweep angle.

FIG. 7 graphically illustrates a leading edge sweep angle relative to aspan position for several example airfoils.

FIG. 8 graphically illustrates a trailing edge sweep angle relative to aspan position for several example airfoils.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures. That is, the disclosedairfoils may be used for engine configurations such as, for example,direct fan drives, or two- or three-spool engines with a speed changemechanism coupling the fan with a compressor or a turbine sections.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis X relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisX which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.55. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45. In anothernon-limiting embodiment the low fan pressure ratio is from 1.1 to 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1200ft/second (365.7 meters/second).

Referring to FIG. 2, which schematically illustrates an example lowpressure compressor (LPC) 44, a variable inlet guide vane (IGV) isarranged downstream from a fan exit stator (FES). The figure is highlyschematic, and the geometry and orientation of various features may beother than shown. An actuator driven by a controller actuates the IGVabout their respective axes. Multiple airfoils are arranged downstreamfrom the IGV. The airfoils include alternating stages of rotors (ROTOR1,ROTOR2, ROTOR3, ROTOR4) and stators (STATOR1, STATOR2, STATOR3,STATOR4). In one example, the low pressure compressor iscounter-rotating relative to the fan blades. In the example shown inFIG. 2, the LPC includes four rotors alternating with four stators.However, in another example, a different number of rotors and statorsmay be used. Moreover, the IGV and stator stages may all be variable,fixed or a combination thereof.

The disclosed airfoils may be used in a low pressure compressor of a twospool engine or in portions of other compressor configurations, such aslow, intermediate and/or high pressure areas of a three spool engine.However, it should be understood that any of the disclosed airfoils maybe used for blades or vanes, and in any of the compressor section,turbine section and fan section.

Referring to FIG. 3, span positions on an airfoil 64 are schematicallyillustrated from 0% to 100% in 10% increments. Each section at a givenspan position is provided by a conical cut that corresponds to the shapeof the core flow path, as shown by the large dashed lines. In the caseof an airfoil with an integral platform, the 0% span positioncorresponds to the radially innermost location where the airfoil meetsthe fillet joining the airfoil to the inner platform. In the case of anairfoil without an integral platform, the 0% span position correspondsto the radially innermost location where the discrete platform meets theexterior surface of the airfoil. For airfoils having no outer platform,such as blades, the 100% span position corresponds to the tip 66. Forairfoils having no platform at the inner diameter, such as cantileveredstators, the 0% span position corresponds to the inner diameter locationof the airfoil. For stators, the 100% span position corresponds to theoutermost location where the airfoil meets the fillet joining theairfoil to the outer platform.

Airfoils in each stage of the LPC are specifically designed radiallyfrom an inner airfoil location (0% span) to an outer airfoil location(100% span) and along circumferentially opposite pressure and suctionsides 72, 74 extending in chord between a leading and trailing edges 68,70 (see FIG. 4). Each airfoil is specifically twisted with acorresponding stagger angle and bent with specific sweep and/or dihedralangles along the airfoil. Airfoil geometric shapes, stacking offsets,chord profiles, stagger angles, sweep and dihedral angles, among otherassociated features, are incorporated individually or collectively toimprove characteristics such as aerodynamic efficiency, structuralintegrity, and vibration mitigation, for example, in a gas turbineengine with a geared architecture in view of the higher LPC rotationalspeeds.

The airfoil 64 has an exterior surface 76 providing a contour thatextends from a leading edge 68 generally aftward in a chord-wisedirection H to a trailing edge 70, as shown in FIG. 4. Pressure andsuction sides 72, 74 join one another at the leading and trailing edges68, 70 and are spaced apart from one another in an airfoil thicknessdirection T. An array of airfoils 64 are positioned about the axis X(corresponding to an X direction) in a circumferential or tangentialdirection Y. Any suitable number of airfoils may be used for aparticular stage in a given engine application.

The exterior surface 76 of the airfoil 64 generates lift based upon itsgeometry and directs flow along the core flow path C. The airfoil 64 maybe constructed from a composite material, or an aluminum alloy ortitanium alloy, or a combination of one or more of these.Abrasion-resistant coatings or other protective coatings may be appliedto the airfoil. The rotor stages may constructed as an integrally bladedrotor, if desired, or discrete blades having roots secured withincorresponding rotor slots of a disc. The stators may be provided byindividual vanes, clusters of vanes, or a full ring of vanes.

Airfoil geometries can be described with respect to various parametersprovided. The disclosed graph(s) illustrate the relationships betweenthe referenced parameters within 10% of the desired values, whichcorrespond to a hot aerodynamic design point for the airfoil. In anotherexample, the referenced parameters are within 5% of the desired values,and in another example, the reference parameters are within 2% of thedesired values. It should be understood that the airfoils may beoriented differently than depicted, depending on the rotationaldirection of the blades. The signs (positive or negative) used, if any,in the graphs of this disclosure are controlling and the drawings shouldthen be understood as a schematic representation of one example airfoilif inconsistent with the graphs. The signs in this disclosure, includingany graphs, comply with the “right hand rule.” The leading and trailingedge sweep angles vary with position along the span, and varies betweena hot, running condition and a cold, static (“on the bench”) condition.

The geared architecture 48 of the disclosed example permits the fan 42to be driven by the low pressure turbine 46 through the low speed spool30 at a lower angular speed than the low pressure turbine 46, whichenables the LPC 44 to rotate at higher, more useful speeds. The leadingand trailing edge sweep angles in a hot, running condition along thespan of the airfoils 64 provides necessary compressor operation incruise at the higher speeds enabled by the geared architecture 48, tothereby enhance aerodynamic functionality and thermal efficiency. Asused herein, the hot, running condition is the condition during cruiseof the gas turbine engine 20. For example, the leading and trailing edgesweep angles in the hot, running condition can be determined in a knownmanner using finite element analysis.

The axial velocity Vx (FIG. 4) of the core flow C is substantiallyconstant across the radius of the flowpath. However the linear velocityU of a rotating airfoil increases with increasing radius. Accordingly,the relative velocity Vr of the working medium at the airfoil leadingedge increases with increasing radius, and at high enough rotationalspeeds, the airfoil experiences supersonic working medium flowvelocities in the vicinity of its tip. The relative velocity at theleading edge 68 is indicated as Vr_(LE), and the relative velocity atthe trailing edge 70 is indicated as Vr_(TE).

Supersonic flow over an airfoil, while beneficial for maximizing thepressurization of the working medium, has the undesirable effect ofreducing fan efficiency by introducing losses in the working medium'stotal pressure. Therefore, it is typical to sweep the airfoil's leadingedge over at least a portion of the blade span so that the workingmedium velocity component in the chordwise direction (perpendicular tothe leading edge) is subsonic. Since the relative velocity Vr increaseswith increasing radius, the sweep angle typically increases withincreasing radius as well. As shown in FIGS. 5 and 6, the sweep angle σat any arbitrary radius Rd (FIG. 3) at the leading edge 68 is indicatedas σ_(LE), and at the trailing edge 70, σ_(TE).

Referring to FIG. 5, the leading edge sweep angle σ_(LE) is the acuteangle between a line 90 tangent to the leading edge 68 of the airfoil 64and a plane 92 perpendicular to the relative velocity vector Vr_(LE).The sweep angle is measured in plane 94, which contains both therelative velocity vector Vr_(LE) and the tangent line 90 and isperpendicular to plane 92. FIG. 7 is provided in conformance with thisdefinition of the leading edge sweep angle σ_(LE).

Referring to FIG. 6, the trailing edge sweep angle σ_(TE) is the acuteangle between a line 96 tangent to the trailing edge 70 of the airfoil64 and a plane 98 perpendicular to the relative velocity vector Vr_(TE).The sweep angle is measured in plane 100, which contains both therelative velocity vector Vr_(TE) and the tangent line 96 and isperpendicular to plane 98. FIG. 8 is provided in conformance with thisdefinition of the trailing edge sweep angle σ_(TE).

A positive X value corresponds to the aftward direction along theengine's axis of rotation. A negative X value corresponds to the forwarddirection along the engine's axis of rotation. Thus, a negative sweepangle indicates an airfoil edge that is oriented in a direction oppositethat of the velocity vector (Vr_(LE) or Vr_(TE)), and a positive sweepangle indicates an airfoil edge that is oriented in the same directionas the velocity vector (Vr_(LE) or Vr_(TE)).

Example relationships between the leading edge sweep (LE SWEEP^(∘)) andthe span position (LE SPAN %) are shown in FIG. 7 for several exampleairfoils, each represented by a curve 110, 112, 114, 116. The airfoilshave a relationship between leading edge sweep angle and span positiondefined by a curve in which the leading edge sweep angle is positive at0% span and crosses to a negative leading edge sweep angle at a spanposition less than 80% span, In one example, the curves cross to thenegative leading edge sweep angle at a span position less than 75% span.In another example, the curves 112, 114 cross to the negative leadingedge sweep angle in a range of 60% span to less than 75% span, and, forexample, the curve 110 crosses to the negative leading edge sweep angleat a span position less than 40% span.

Example relationships between the trailing edge sweep (TE SWEEP^(∘)) andthe span position (TE SPAN %) are shown in FIG. 8 for several exampleairfoils, each represented by a curve 118, 120, 122, 124. In oneexample, the airfoils have a relationship between trailing edge sweepangle and span position defined by curves 118, 120 in which the trailingedge sweep angle is negative at 0% span and changes less than 10° from0% span to 90% span. In another example, the airfoils have arelationship between trailing edge sweep angle and span position definedby curves 122, 124 in which the trailing edge sweep angle is positive at0% span and has a decrease in trailing edge sweep angle of greater than10° from 0% span to 40% span.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A compressor airfoil of a turbine enginecomprising: pressure and suction sides extending in a radial directionfrom a 0% span position to a 100% span position, wherein the airfoil hasa relationship between leading edge sweep angle and span positiondefined by a curve in which the leading edge sweep angle is positive at0% span and crosses to a negative leading edge sweep angle at a spanposition less than 80% span, wherein a negative sweep angle is in theforward direction, and a positive sweep angle is in the rearwarddirection, wherein the airfoil has a relationship between trailing edgesweep angle and span position defined by a curve in which the trailingedge sweep angle is negative at 0% span and changes less than 10° from0% span to 90% span.
 2. A compressor airfoil of a turbine enginecomprising: pressure and suction sides extending in a radial directionfrom a 0% span position to a 100% span position, wherein the airfoil hasa relationship between leading edge sweep angle and span positiondefined by a curve in which the leading edge sweep angle is positive at0% span and crosses to a negative leading edge sweep angle at a spanposition less than 80% span, wherein a negative sweep angle is in theforward direction, and a positive sweep angle is in the rearwarddirection, wherein the airfoil has a relationship between trailing edgesweep angle and span position defined by a curve in which the trailingedge sweep angle is positive at 0% span and has a decrease in trailingedge sweep angle of greater than 10° from 0% span to 40% span, whereinthe trailing edge sweep angle is strictly decreasing from 0% span to 40%span.
 3. The compressor airfoil according to claim 2, wherein the curvecrosses to the negative leading edge sweep angle at a span position lessthan 75% span.
 4. The compressor airfoil according to claim 3, whereinthe curve crosses to the negative leading edge sweep angle in a range of60% span to less than 75% span.
 5. The compressor airfoil according toclaim 2, wherein the curve crosses to the negative leading edge sweepangle at a span position less than 40% span.
 6. The compressor airfoilaccording to claim 2, wherein the positive trailing edge sweep angleindicates an airfoil edge that is oriented in the same direction as thevelocity vector of an incoming gas flow relative to the leading ortrailing edge of the airfoil.
 7. A gas turbine engine comprising: acombustor section arranged between a compressor section and a turbinesection, wherein the compressor section includes at least a low pressurecompressor and a high pressure compressor, the high pressure compressorarranged immediately upstream of the combustor section; a fan sectionhaving an array of twenty-six or fewer fan blades, wherein the lowpressure compressor is immediately downstream from the fan section,wherein the low pressure compressor is counter-rotating relative to thefan blades; and an airfoil arranged in the compressor section outsidethe high pressure compressor, and the airfoil includes pressure andsuction sides extending in a radial direction from a 0% span position atan inner flow path location to a 100% span position at an airfoil tip,wherein the airfoil has a relationship between leading edge sweep angleand span position defined by a curve in which the leading edge sweepangle is positive at 0% span and crosses to a negative leading edgesweep angle at a span position less than 80% span, wherein a negativeleading edge sweep angle is in the forward direction, and a positiveleading edge sweep angle is in the rearward direction.
 8. The gasturbine engine according to claim 7, wherein the gas turbine engine is atwo-spool configuration.
 9. The gas turbine engine according to claim 7,wherein the airfoil is rotatable relative to an engine axis of an enginestatic structure.
 10. The gas turbine engine according to claim 7,wherein the curve crosses to the negative leading edge sweep angle at aspan position less than 75% span.
 11. The gas turbine engine accordingto claim 10, wherein the curve crosses to the negative leading edgesweep angle in a range of 60% span to less than 75% span.
 12. The gasturbine engine according to claim 10, wherein the curve crosses to thenegative leading edge sweep angle at a span position less than 40% span.13. The gas turbine engine according to claim 7, wherein the airfoil hasa relationship between trailing edge sweep angle and span positiondefined by a curve in which the trailing edge sweep angle is negative at0% span and changes less than 10° from 0% span to 90% span.
 14. The gasturbine engine according to claim 7, wherein the airfoil has arelationship between trailing edge sweep angle and span position definedby a curve in which the trailing edge sweep angle is positive at 0% spanand has a decrease in trailing edge sweep angle of greater than 10° from0% span to 40% span.
 15. The gas turbine engine according to claim 7,wherein the positive leading edge sweep angle is in the aftwarddirection along the gas turbine engine's axis of rotation.